The present invention relates generally to aircraft gas turbine engines with cooling cavities adjacent an exhaust flowpath of the engine downstream of an afterburner and, more specifically, downstream of an afterburner with a trapped vortex cavity.
High performance military aircraft typically include a turbofan gas turbine engine having an afterburner or augmentor for providing additional thrust when desired particularly for supersonic flight. The turbofan engine includes in downstream serial flow communication, a multistage fan, a multistage compressor, a combustor, a high pressure turbine powering the compressor, and a low pressure turbine powering the fan. A bypass duct surrounds and allows a portion of the fan air to bypass the multistage compressor, combustor, high pressure, and low pressure turbine.
During operation, air is compressed in turn through the fan and compressor and mixed with fuel in the combustor and ignited for generating hot combustion gases which flow downstream through the turbine stages which extract energy therefrom. The hot core gases are then discharged into an exhaust section of the engine which includes an afterburner from which they are discharged from the engine through a variable area exhaust nozzle.
Afterburners are located in exhaust sections of engines which includes an exhaust casing and an exhaust liner circumscribing a combustion zone. Fuel injectors (such as spraybars) and flameholders are mounted between the turbines and the exhaust nozzle for injecting additional fuel when desired during reheat operation for burning in the afterburner for producing additional thrust. Thrust augmentation or reheat using such fuel injection is referred to as wet operation while operating dry refers to not using the thrust augmentation. The annular bypass duct extends from the fan to the afterburner for bypassing a portion of the fan air around the core engine to the afterburner. This bypass air is mixed with the core gases and fuel from the spraybars, ignited, and combusted prior to discharge through the exhaust nozzle. The bypass air is also used in part for cooling the exhaust liner.
Various types of flameholders are known and provide local low velocity recirculation and stagnation regions therebehind, in regions of otherwise high velocity core gases, for sustaining and stabilizing combustion during reheat operation. Since the core gases are the product of combustion in the core engine, they are initially hot, and are further heated when burned with the bypass air and additional fuel during reheat operation. Augmentors currently are used to maximize thrust increases and tend to be full stream and consume all available oxygen in the combustion process yielding high augmentation ratios for example about 70%.
In regions immediately downstream of the flameholder, the gas flow is partially recirculated and the velocity is less than the rate of flame propagation. In these regions, there will be a stable flame existing which can ignite new fuel as it passes. Unfortunately, flameholders in the gas stream inherently cause flow losses and reduced engine efficiency. Several modern gas turbine engines and designs include radially extending spraybars and flameholders in an effort to improve flame stability and reduce the flow losses. Radial spraybars integrated with radial flameholders are disclosed in U.S. Pat. Nos. 5,396,763 and 5,813,221. Radial spraybars disposed between radial flameholders having integrated radial spraybars have been incorporated in the GE F414 and GE F110-132 aircraft gas turbine engines. This arrangement provides additional dispersion of the fuel for more efficient combustion and unload fueling of the radial flameholders with the integrated radial spraybars so that they do not blowout and/or have unstable combustion due to excess fueling.
Since fuel is typically injected upstream of the flameholders, undesirable auto-ignition of the fuel and combustion which might occur upstream of the flameholders causes flameholder distress which also significantly reduces the useful life of the flameholders. Since V-gutter flameholders are suspended within the core gases, they are more difficult to effectively cool and, typically, experience circumferential variation in temperature, which correspondingly effects thermal stress, which also decreases the useful life thereof. V-gutter flameholders have limited flameholding capability and their aerodynamic performance and characteristics negatively impact the size, performance, and thrust capability of the engine. This is, in part, due to the combustion zone having sufficient length to allow substantially complete combustion of the fuel added by the spraybars prior to discharge through the nozzle and wide ranging flight speeds and Mach numbers. Flame stabilizing trapped vortex cavity pilots have been developed to provide better performance characteristics than previous afterburners or augmentors with lower flow losses and improved engine efficiency. Internally and externally fueled annular trapped vortex cavity pilots have a cavity opening open to an exhaust flowpath. See U.S. Pat. No. 7,225,623 entitled “Trapped Vortex Cavity Afterburner”, U.S. Pat. No. 7,467,518 entitled “Externally Fueled Trapped Vortex Cavity Augmentor”, and US Patent Publication No. 2009/0056340A1 entitled “Augmentor With Trapped Vortex Cavity Pilot”, now U.S. Pat. No. 8,011,188.
Between the exhaust nozzle and the trapped vortex cavity pilot is the exhaust casing and exhaust liner circumscribing the combustion zone. The trapped vortex cavity may be formed in a trapped vortex liner upstream and separate from the exhaust liner, thus, forming a cavity between the two liners. It is important to purge this cavity because it can cause flow separation and therefore create a flame-holding region and burning of the liners in this area which can lead to premature wear or failure and can pose a flight safety hazard. Thus, it is highly desirable to provide a good cavity purge to prevent burning in the area of the cavity.